Rotor keyhole fillet for a gas turbine engine

ABSTRACT

A rotor for a gas turbine engine includes an annular structure having a blade slot. A hub engagement feature is provided on the annular structure. The hub engagement feature includes first and second surfaces transverse to one another and joined by a fillet that is recessed with respect to the first and second surfaces. A method of manufacturing a rotor includes the steps of machining an annular hub engagement feature into a rotor. The hub engagement feature includes first and second surfaces transverse to one another and is joined by a fillet that is recessed with respect to the first and second surfaces. The method includes the step of peening the fillet, and grinding the first and second surfaces.

BACKGROUND

This disclosure relates to a rotor for a gas turbine engine. Moreparticularly, the disclosure relates to fillet geometry for the rotor.

A typical gas turbine engine includes multiple of compressor stagesupstream from a combustor section. A turbine section is arrangeddownstream from the combustor. In one example configuration, at leastone end of a compressor rotor is secured to a shaft by a hub. The hubengages a hub engagement feature on the rotor to secure the rotorrelative to the shaft. Typically, a nut is received on thecorrespondingly threaded portion of the shaft and applies a clampingload to the rotor via the hub.

The hub engagement feature on the rotor is provided by first and secondannular surfaces that are at a right angle to one another. A filletjoins the first and second surfaces, which are arranged tangentiallyrelative to the fillet.

SUMMARY

In one exemplary embodiment, a rotor for a gas turbine engine includesan annular structure having a blade slot. A hub engagement feature isprovided on the annular structure. The hub engagement feature includesfirst and second surfaces transverse to one another and joined by afillet that is recessed with respect to the first and second surfaces.

In a further embodiment of any of the above, the first and secondsurfaces are normal to one another.

In a further embodiment of any of the above, the fillet includes apeened surface.

In a further embodiment of any of the above, the annular structure isconstructed from a nickel alloy.

In a further embodiment of any of the above, the first and secondsurfaces are provided by ground surfaces.

In a further embodiment of any of the above, the first and secondsurfaces are non-tangential to the fillet.

In another exemplary embodiment, a rotor assembly for a gas turbineengine includes a shaft. The rotor assembly includes a rotor thatsupports a blade and includes a hub engagement feature. The hubengagement feature includes first and second rotor surfaces transverseto one another and is joined by a fillet that is recessed with respectto the first and second rotor surfaces. A hub is supported on the shaftand engages the hub engagement feature.

In a further embodiment of any of the above, the rotor assembly includesa nut secured to the shaft and applies a clamping load to the hubengagement feature via the hub.

In a further embodiment of any of the above, the hub includes first andsecond hub surfaces respectively engaging the first and second rotorsurfaces under the clamping load. The first and second hub surfaces arespaced from the fillet.

In a further embodiment of any of the above, the first and second rotorsurfaces are normal to one another.

In a further embodiment of any of the above, the fillet includes apeened surface.

In a further embodiment of any of the above, the rotor is constructedfrom a nickel alloy.

In a further embodiment of any of the above, the first and second rotorsurfaces are provided by ground surfaces.

In a further embodiment of any of the above, the first and second rotorsurfaces are non-tangential to the fillet.

In another exemplary embodiment, a method of manufacturing a rotorincludes the steps of machining an annular hub engagement feature into arotor. The hub engagement feature includes first and second surfacestransverse to one another and is joined by a fillet that is recessedwith respect to the first and second surfaces. The method includes thestep of peening the fillet, and grinding the first and second surfaces.

In a further embodiment of any of the above, the grinding step isperformed after the peening step.

In a further embodiment of any of the above, the first and secondsurfaces are non-tangential to the fillet.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 is a schematic cross-sectional view of an example gas turbineengine.

FIG. 2 is a schematic view of an example rotor and hub supported by ashaft.

FIG. 3 is an enlarged cross-sectional view of a portion of the rotor andhub.

FIG. 4 is a flow chart depicting an example method of manufacturing arotor.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example gas turbine engine 20 thatincludes a fan section 22, a compressor section 24, a combustor section26 and a turbine section 28. Alternative engines might include anaugmenter section (not shown) among other systems or features. The fansection 22 drives air along a bypass flow path B while the compressorsection 24 draws air in along a core flow path C where air is compressedand communicated to a combustor section 26. In the combustor section 26,air is mixed with fuel and ignited to generate a high pressure exhaustgas stream that expands through the turbine section 28 where energy isextracted and utilized to drive the fan section 22 and the compressorsection 24.

Although the disclosed non-limiting embodiment depicts a turbofan gasturbine engine, it should be understood that the concepts describedherein are not limited to use with turbofans as the teachings may beapplied to other types of turbine engines; for example a turbine engineincluding a three-spool architecture in which three spoolsconcentrically rotate about a common axis and where a low spool enablesa low pressure turbine to drive a fan via a gearbox, an intermediatespool that enables an intermediate pressure turbine to drive a firstcompressor of the compressor section, and a high spool that enables ahigh pressure turbine to drive a high pressure compressor of thecompressor section.

The example engine 20 generally includes a low speed spool 30 and a highspeed spool 32 mounted for rotation about an engine central longitudinalaxis A relative to an engine static structure 36 via several bearingsystems 38. It should be understood that various bearing systems 38 atvarious locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatconnects a fan 42 and a low pressure (or first) compressor section 44 toa low pressure (or first) turbine section 46. The inner shaft 40 drivesthe fan 42 through a speed change device, such as a geared architecture48, to drive the fan 42 at a lower speed than the low speed spool 30.The high-speed spool 32 includes an outer shaft 50 that interconnects ahigh pressure (or second) compressor section 52 and a high pressure (orsecond) turbine section 54. The inner shaft 40 and the outer shaft 50are concentric and rotate via the bearing systems 38 about the enginecentral longitudinal axis A.

A combustor 56 is arranged between the high pressure compressor 52 andthe high pressure turbine 54. In one example, the high pressure turbine54 includes at least two stages to provide a double stage high pressureturbine 54. In another example, the high pressure turbine 54 includesonly a single stage. As used herein, a “high pressure” compressor orturbine experiences a higher pressure than a corresponding “lowpressure” compressor or turbine.

The core airflow C is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a star gear systemor other gear system, with a gear reduction ratio of greater than about2.3 and the low pressure turbine 46 has a pressure ratio that is greaterthan about 5. In one disclosed embodiment, the engine 20 bypass ratio isgreater than about ten (10:1), the fan diameter is significantly largerthan that of the low pressure compressor 44, and the low pressureturbine 46 has a pressure ratio that is greater than about 5:1. Lowpressure turbine 46 pressure ratio is pressure measured prior to inletof low pressure turbine 46 as related to the pressure at the outlet ofthe low pressure turbine 46 prior to an exhaust nozzle. It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentinvention is applicable to other gas turbine engines including directdrive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned per hour divided by lbf of thrustthe engine produces at that minimum point. “Fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tambient degR)/518.7)̂0.5]. The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

Referring to FIG. 2, the high pressure compressor 52 includes a rotor 60supported relative to the outer shaft 50. In an example, the rotor 60 isthe last stage disk of the high pressure compressor 52. The hub 62supports the rotor 60 relative to the outer shaft 52. The hub 62 issecured to the outer shaft 50 by a nut 70, which applies a clampingforce to a hub engagement feature 71 (discussed below) on the rotor 62.

The rotor 60, which is constructed from a nickel alloy, includes one ormore slots 65 that support multiple circumferentially spaced blades 64.It should be understood, however, that the blades 64 may be integratedwith the rotor 60.

A first air seal 66 is supported by the rotor 60, in the example, whichcooperates with a second air seal 68 supported by the engine staticstructure 36. The first air seal 66 may integral with or separate fromthe rotor 62.

The hub engagement feature 71 is shown in more detail in FIG. 3. Firstand second rotor surfaces 72, 74 are provided on the rotor 60 normal(that is, perpendicular) to one another. The hub 62 includes first andsecond hub surfaces 76, 78 that also are normal with respect to oneanother. The first and second hub surfaces 76, 78 respectively engagethe first and second rotor surfaces 72, 74 of the hub engagement feature71. A fillet 80 joins the first and second rotor surfaces 72, 74 to oneanother. The fillet 80 is recessed with respect to the first and secondrotor surfaces 72, 74 such that the first and second rotor surfaces 72,74 are not tangential to the fillet 80. The first and second rotorsurfaces 72, 74 meet the fillet 80 at first and second intersections 82,84, respectively, providing a keyhole-shaped fillet geometry. Saidanother way, the first and second rotor surfaces 72, 74 are proud of, orextend above, the fillet 80. The first and second rotor surfaces 72, 74extend toward the fillet 80 beyond the first and second hub surfaces 76,78.

Referring to FIG. 4, a method 86 of manufacturing the rotor 60 isdescribed. The annular hub engagement feature 71 is machined into therotor 60, as indicated at block 88. The hub engagement feature 71includes the first and second rotor surfaces 72, 74, which aretransverse to one another, for example, at a right angle. The fillet 80may be machined at the same time as the first and second rotor surfaces72, 74, subsequently to the first and second rotor surfaces 72, 74, orprior to forming the first and second rotor surfaces 72, 74. The fillet80 is peened to relieve the stresses in this typically high stressedarea, as indicated at block 90. The peening operation may occur beforeor after machining of the first and second rotor surfaces 72, 74. Thefirst and second rotor surfaces 72, 74 may receive a finish grinding, asindicated at block 92, to provide a dimensionally precise surface withdesired surface finish for receiving the hub 62.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. A rotor for a gas turbine engine comprising: anannular structure having a blade slot; and a hub engagement featureprovided on the annular structure, the hub engagement feature includingfirst and second surfaces transverse to one another and joined by afillet that is recessed with respect to the first and second surfaces.2. The rotor according to claim 1, wherein the first and second surfacesare normal to one another.
 3. The rotor according to claim 1, whereinthe fillet includes a peened surface.
 4. The rotor according to claim 3,wherein the annular structure is constructed from a nickel alloy.
 5. Therotor according to claim 3, wherein the first and second surfaces areprovided by ground surfaces.
 6. The rotor according to claim 1, whereinthe first and second surfaces are non-tangential to the fillet.
 7. Arotor assembly for a gas turbine engine comprising: a shaft; a rotorsupporting a blade and including a hub engagement feature, the hubengagement feature including first and second rotor surfaces transverseto one another and joined by a fillet that is recessed with respect tothe first and second rotor surfaces; and a hub supported on the shaftand engaging the hub engagement feature.
 8. The rotor assembly accordingto claim 7, comprising a nut secured to the shaft and applying aclamping load to the hub engagement feature via the hub.
 9. The rotorassembly according to claim 8, wherein the hub includes first and secondhub surfaces respectively engaging the first and second rotor surfacesunder the clamping load, the first and second hub surfaces spaced fromthe fillet.
 10. The rotor assembly according to claim 9, wherein thefirst and second rotor surfaces are normal to one another.
 11. The rotorassembly according to claim 7, wherein the fillet includes a peenedsurface.
 12. The rotor assembly according to claim 11, wherein the rotoris constructed from a nickel alloy.
 13. The rotor assembly according toclaim 11, wherein the first and second rotor surfaces are provided byground surfaces.
 14. The rotor assembly according to claim 7, whereinthe first and second rotor surfaces are non-tangential to the fillet.15. A method of manufacturing a rotor comprising the steps of: machiningan annular hub engagement feature into a rotor, the hub engagementfeature including first and second surfaces transverse to one anotherand joined by a fillet that is recessed with respect to the first andsecond surfaces; peening the fillet; and grinding the first and secondsurfaces.
 16. The method according to claim 15, wherein the grindingstep is performed after the peening step.
 17. The method according toclaim 15, wherein the first and second surfaces are non-tangential tothe fillet.